Historically, gas turbine engines have been able to achieve much better life and reliability than internal combustion (IC) engines, perhaps because they employ a steady flow combustion process, fewer moving parts, no reciprocating motion, and no surfaces in rubbing contact. Gas turbines are typically also adaptable to using a wider range of fuels. They generate almost no vibration and steady output torque, while IC engines more commonly generate peak:mean torque levels between 4:1 and 16:1. Turbines also produce only high-frequency noise, which is easier to mitigate and tends to naturally damp out quickly in the atmosphere, while low sound frequencies can travel long distances through the air, ground, or water. With all these advantages, gas turbines are often the preferred power plant in many applications, particularly aircraft propulsion and stationary power generation.
Small gas turbines are of interest for residential combined heat and power (CHP), unmanned air vehicle (UAV) propulsion, portable power generation, and other applications. However, they have difficulty competing with piston engines due to inferior fuel efficiencies. This may be due to scaling effects. For example, gas flows at small length scales are characterized by low Reynolds numbers, which means that inertial effects become less important than viscous effects, viscous friction wastes more power, and turbomachinery becomes less efficient. Small turbomachinery tip clearances and trailing edge thicknesses tend to be relatively larger, relative to blade chord and span, than in large engines, and this also leads to larger losses and lower compression and expansion efficiencies. Another scaling effect is heat transfer by conduction and convection, which tends to be more effective at small scale than at large scale. Unwanted heat transfers, particularly heat losses from the burner and heat transfer to the compressor, can reduce fuel efficiency severely in small turbines.
However, scaling effects that favor small engines can be exploited to improve their fuel efficiency. For example, the fact that convection heat transfer improves at small scale suggests that a “recuperator”—a heat exchanger that preheats combustion air using waste heat from the exhaust, saving fuel—could perform well in a small engine. Ceramic materials are more reliable at small scale, and since they can withstand higher temperatures than metals without cooling, they enable higher turbine inlet temperatures, which improve engine efficiency and specific work. Higher specific work means the engine uses less airflow to produce its target power output. This reduces recuperator size. Recuperators can be made from ceramic materials as well, which reduces weight and improves performance.
One early effort to implement the recuperation strategy was the IHI Dynajet, a commercial three kilowatt recuperated gas turbine designed for portable power generation. It was designed for 20% fuel efficiency, and its compressor, turbine, and recuperator performed well enough in separate rig tests to achieve this goal. However, when these components were assembled into a complete engine, it only reached 6% efficiency. It was shown that fluid leaks among the components and heat leakage from the hot section into the compressor flow path were the main reasons for the discrepancy, and that many of these flows could be greatly reduced through better thermal insulation and gas seals.
Additional efforts involved companies developing small turbine engines or “microturbines” for small-scale power generation in the 30-200 kW power range, with efficiencies well above 25%. These small turbine engines are typically manufactured and sold in large quantities for markets like combined heat and power for small to medium businesses, backup power for hospitals, and natural gas propulsion for commercial vehicles. Others have sought to build smaller engines to fit the CHP demands of a single residence. Since these engines are designed for land-based applications where weight does not matter as much as fuel efficiency, a relatively large and heavy recuperator could be used, and some complexity could be tolerated. These engines typically employ a conventional turbomachinery arrangement in which the compressor outlet and turbine inlet are adjacent to each other.
FIG. 1 is a prior art diagram of a conventional recuperated microturbine. The combustor is an annular design surrounding the turbine and its exhaust duct. Annular combustors are common in gas turbines but they are somewhat complex, requiring multiple fuel injectors to be spaced evenly around the ring, to spread out the heat of combustion evenly.
In some applications like aircraft propulsion, compactness and weight carry greater importance. Simplicity is even more critical in aircraft, partly because simpler engines tend to be smaller, lighter, and cheaper, but more importantly because simplicity directly leads to better reliability. One way to simplify the engine is to replace the annular combustor with a single “can” combustor and just one fuel injector.
One of the most successful turboshaft engines of all time, the Rolls Royce Allison 250 and its many derivatives, accomplished this by reversing the turbine gas flow direction, so that the turbine outlet faces the compressor. The inlet faces away from the rotating assembly, where a can combustor can be located; it need not be annular because there is no turbomachinery or exhaust duct occupying its centerline. This innovation simplifies the whole engine dramatically, while also making it much easier to inspect and repair. To make use of these advantages in recuperated engines, others designed gas turbine engines with reversed turbines and simple can combustors. FIG. 2 is a prior art diagram of a gas turbine engine with a reversed turbine. This configuration has been shown to be convenient and compact to locate the recuperator in the space surrounding the burner. The fuel injector, not shown in FIG. 2, can typically be positioned on the left side of the picture on the engine centerline.
The layout of the engine in the prior art FIG. 2 is compact and simple, and it can reduce heat losses from the combustor. A single can combustor typically has much less surface area than an annular combustor of the same volume. This is partly because it has only an outer surface (“liner”), rather than an outer and an inner liner; and because the outer diameter of the annular combustor must be larger, to provide the same volume in the same axial length. With much less surface area, heat transfer by convection and radiation to the surroundings will typically be less, and with smaller heat losses, less fuel is required to heat up the gas to the target combustor outlet temperature. Thus the can combustor, with its reduced surface area, will typically be more fuel-efficient.
Although advantageous in many respects, the FIG. 2 layout does suffer from certain drawbacks. One problem is the axial thrust on the bearings. Compressor and turbine rotors generate thrust due to imbalanced gas pressures. The back (flat) face of a compressor wheel is exposed to elevated gas pressure, roughly equal to that at the rotor outlet/diffuser (compressor stator) inlet, while gas pressures on the front (inlet) side of the rotor are lower. Thus, there is a net axial thrust toward the inlet. In turbine rotors, gas pressure is higher on the inlet side than it is on the outlet (exhaust) side. Therefore, in FIG. 2, both the turbine and the compressor generate an axial thrust on the rotating assembly that points from left to right. This means that their values add together rather than balancing each other out, leading to potentially very high net axial loads on the bearings. Since the life of a rolling element bearing is proportional to the load cubed, this can lead to short bearing life and premature bearing failure.
Along with high bearing thrust loads, another disadvantage of the FIG. 2 prior art layout is gas leakage. With the compressor back face (where gas pressure is high) located next to the turbine outlet (where gas pressure is low), the leakage rate of gas from compressor to turbine is likely to be very high. To minimize this, there would need to be a gas seal positioned somewhere along the rotating shaft between the turbine and compressor rotors. However, this shaft will be hot. It would be difficult to position a contact seal in that location because the contact seal would overheat. A non-contact labyrinth seal might make sense but due to the high gas pressure difference, the leakage rate would be quite substantial through this area, and this would substantially reduce the overall engine efficiency. Based on published studies on small gas turbine engines, it is very important to minimize leakage flows in all areas of the engine. In this engine layout there is no readily apparent way to minimize the gas leakage rate along the rotating shaft.
Another disadvantage of the engine layouts in both prior art FIG. 1 and FIG. 2 relates to heat transfer to the compressor diffuser. The diffuser is the radial channel from the compressor rotor outlet to the outer engine casing. In the diffuser, air enters at high tangential velocity and slows down as it progresses toward the diffuser outlet. During that process, its total pressure (static+dynamic) remains relatively constant, falling only slightly due to friction, but as the air slows down, the dynamic pressure decreases, while the static pressure—the useful part—rises. In fact, a substantial portion of the overall compressor pressure rise occurs in the diffuser, often approximately 40%. The warmer the air is while it is slowing down and converting dynamic pressure to static pressure, the less the static pressure will rise. To make up for this, the rotor has to impart more tangential velocity to the air, either by spinning faster or by being larger in diameter. Either way makes the compressor consume more shaft power and makes the engine, as a whole, less efficient. In other words, heat transfer to the compressor diffuser is bad because it requires more shaft power to compress the gas enough to reach the target outlet pressure. Another way of looking at this is to say that heat transfer to any part of the compressor will reduce its efficiency, relative to an adiabatic (i.e., unheated) compressor. The cooler any gas is during a compression process, the less energy that process takes. In any engine, the cooler the gas is during compression and the hotter it is during expansion, the more net power is generated per unit mass flow, and the higher the efficiency. On a shaft power basis, this effect can be substantial, as many others have shown.
Another problem in many prior art microturbines is the difficulty of keeping the bearings cool, especially the one closest to the turbine rotor. In very small engines, the distance between this bearing and the turbine rotor can be as short as 1-2 inches, and the temperature of the turbine rotor can be 900 degrees C. or more. The bearings must typically operate below 200 degrees C. to avoid damaging the oil and reducing the hardness and load capacity of the steel bearing races. Microturbines sometimes cantilever the turbine and compressor rotors from a pair of bearings located upstream from the compressor inlet (see, e.g., to the right of the compressor rotor in FIG. 2). This keeps them cool, but the cantilevered rotor then does not have room for an alternator on the same shaft. If the rotor assembly is made longer to make room, shaft dynamics problems can result. The engine in FIG. 1 employs air bearings, a good solution that does eliminate the need for oil, a big advantage. However, air bearings have some disadvantages as well, including air foil touchdown during start/stop operation, the need to be kept very clean due to tight clearances, complex rotordynamics, relatively high cost, and the need for considerable testing and development work to ensure high reliability.
Accordingly, there remains a need in the art for a simple, compact, lightweight, efficient recuperated gas turbine engine configuration that reduces gas leakage and hot section heat losses, improves compressor efficiency by reducing heat transfer to the compressor rotor and diffuser, and minimizes axial bearing loads by opposing the turbine and compressor axial thrust directions.